Coating for improved surface finish

ABSTRACT

A coating includes: at least 34.9 percent by mass silicon dioxide; at least 9.1 percent by mass aluminum oxide; and at least 16.1 percent by mass yttrium oxide.

STATEMENT OF GOVERNMENT RIGHTS

The invention described in the present disclosure was made with thesupport of the U.S. Government under contract number DE-FE0024006, whichwas awarded by the Department of Energy. The U.S. Government has certainrights in this invention.

BACKGROUND

The present subject matter relates generally to coatings, and morespecifically to coatings for gas turbine engines.

As the demand for more efficient gas turbine engines drives internaloperating temperatures higher, the transition to higher temperaturematerials has driven the transition from metallic nickel-basedsuperalloys to ceramic matrix composites (CMC), which offer bothmechanical strength as well as resistance to high temperatures. However,CMCs may have higher production costs and/or longer manufacturing cycletimes compared to metallic nickel-based superalloys. In addition, thegeometries and/or topologies in which CMC components can be formed maybe limited compared to metallic nickel-based superalloys.

The orthotropic nature of CMCs can lead to porosity and inconsistentmachined surface finish. Porosity in a CMC sealing surfaces (i.e., inapplications where CMCs are installed within or proximate gas turbinehot gas paths (HGP)) may allow cooling flow leakage past the seals,thereby resulting in efficiency losses. In addition, inconsistentmachining, due to, for example varying speeds of material removal ofceramic matrices and/or fibers, can lead to ripples or rough surfaces infinished CMC components. Rough surfaces may create undulations and gapsbetween the CMC seals and sealing surfaces for cooling flow to leakthrough, again leading to efficiency losses in gas turbine engines.

BRIEF DESCRIPTION OF THE EMBODIMENTS

Aspects of the present embodiments are summarized below. Theseembodiments are not intended to limit the scope of the present claimedembodiments, but rather, these embodiments are intended only to providea brief summary of possible forms of the embodiments. Furthermore, theembodiments may encompass a variety of forms that may be similar to ordifferent from the embodiments set forth below, commensurate with thescope of the claims.

In one aspect, a coating includes: at least 34.9 percent by mass silicondioxide; at least 9.1 percent by mass aluminum oxide; and at least 16.1percent by mass yttrium oxide.

In another aspect, a coating includes: at least 9.8 percent by bariumoxide; at least 5.2 percent by mass aluminum oxide; and at least 40.3percent by mass silicon dioxide.

In another aspect, a coating system includes a S-A-Y materialcomposition including: from about 49 percent to about 59 percent by masssilicon dioxide; from about 13 percent to about 23 percent by massaluminum oxide; and from about 23 percent to about 33 percent by massyttrium oxide. The coating system includes a B-A-S material compositionincluding: from about 15 percent to about 25 percent by mass bariumoxide; from about 8 to about 18 percent by mass aluminum oxide; and fromabout 62 percent to about 72 percent by mass silicon dioxide.

BRIEF DESCRIPTION OF THE DRAWINGS

These and other features, aspects, and advantages of the presentdisclosure will become better understood when the following detaileddescription is read with reference to the accompanying drawings in whichlike characters represent like parts throughout the drawings, wherein:

FIG. 1 is a side schematic representation of a gas turbine engine;

FIG. 2 is a side schematic representation of a turbine airfoil;

FIG. 3 is summary of coating composition test points;

FIG. 4 is summary of coating composition that were selected for furtherstudy; and

FIG. 5 is a method of forming a CMC component with a coating disposedthereon, according to aspects of the present embodiments.

Unless otherwise indicated, the drawings provided herein are meant toillustrate features of embodiments of the disclosure. These features arebelieved to be applicable in a wide variety of systems comprising one ormore embodiments of the disclosure. As such, the drawings are not meantto include all conventional features known by those of ordinary skill inthe art to be required for the practice of the embodiments disclosedherein.

DETAILED DESCRIPTION

In the following specification and the claims, reference will be made toa number of terms, which shall be defined to have the followingmeanings.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

“Optional” or “optionally” means that the subsequently described eventor circumstance may or may not occur, and that the description includesinstances where the event occurs and instances where it does not.

Approximating language, as used herein throughout the specification andclaims, may be applied to modify any quantitative representation thatcould permissibly vary without resulting in a change in the basicfunction to which it is related. Accordingly, a value modified by a termor terms, such as “about” and “substantially”, are not to be limited tothe precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value. Here and throughout the specification andclaims, range limitations may be combined and/or interchanged, suchranges are identified and include all the sub-ranges contained thereinunless context or language indicates otherwise.

As used herein, the term “axial” refers to a direction aligned with acentral axis or shaft of a gas turbine.

As used herein, the term “circumferential” refers to a direction ordirections around (and tangential to) the outer circumference of the gasturbine, or for example the circle defined by the swept area of therotor of the gas turbine. As used herein, the terms “circumferential”and “tangential” may be synonymous.

As used herein, the term “radial” refers to a direction moving outwardlyaway from the central axis of the gas turbine. A “radially inward”direction is aligned toward the central axis moving toward decreasingradii. A “radially outward” direction is aligned away from the centralaxis moving toward increasing radii.

Referring now to the drawings, wherein like numerals refer to likecomponents, FIG. 1 illustrates an example of a gas turbine 10 which mayincorporate various aspects of the embodiments disclosed herein. Asshown, the gas turbine 10 generally includes a compressor section 12having an inlet 14 disposed at an upstream end of the gas turbine 10,and a casing 16 that at least partially surrounds the compressor section12. The gas turbine 10 further includes a combustion section 18 havingat least one combustor 20 downstream from the compressor section 12, anda turbine section 22 downstream from the combustion section 18. Asshown, the combustion section 18 may include a plurality of thecombustors 20. A shaft 24 extends axially through the gas turbine 10.FIG. 1 illustrates the radial 94, axial 92 and circumferentialdirections 90.

Referring still to FIG. 1, the gas turbine 10 may include a transitionpiece 52 disposed between a downstream end of the combustor 20 and anupstream end of the turbine section 22. The combustor 20 may include acombustor liner 54 defining the boundaries of the combustor 20. Thecombustion section 18 may include a plurality of substantiallycylindrical “can-style” combustors 20 circumferentially spaced aroundthe gas turbine 10, in which case the combustor liners 54 may also besubstantially cylindrical. In other embodiments, the combustion section18 may include an annular combustor 20, in which case the combustorliner 54 may include both an inner liner (not shown) defining a radiallyinner boundary of the annular combustor 20, as well as an outer liner(not shown) defining a radially outer boundary of the annular combustor20. The gas turbine 10 may also include a stage one nozzle 56 located inthe turbine section 22 and disposed axially aft of the transition piece52, as well as a second stage nozzle 58 disposed downstream of a firststage turbine rotor 36. The gas turbine 10 may also include one or moreflow path ducts 60 defining a radially outer boundary of a turbine gaspath at axial locations between rotors and stators (i.e., blades 36 andnozzles 34). A gas turbine “hot section” may include both the combustorsection 18 and components thereof, as well as the turbine section 22 andcomponents thereof.

In operation, air 26 is drawn into the inlet 14 of the compressorsection 12 and is progressively compressed to provide compressed air 28to the combustion section 18. The compressed air 28 flows into thecombustion section 18 and is mixed with fuel in the combustor 20 to forma combustible mixture. The combustible mixture is burned in thecombustor 20, thereby generating a hot gas 30 that flows from thecombustor 20 across a first stage 32 of turbine nozzles 34 and into theturbine section 22. The turbine section generally includes one or morerows of rotor blades 36 axially separated by an adjacent row of theturbine nozzles 34. The rotor blades 36 are coupled to the rotor shaft24 via a rotor disk. The rotor shaft 24 rotates about an enginecenterline CL. A turbine casing 38 at least partially encases the rotorblades 36 and the turbine nozzles 34. Each or some of the rows of rotorblades 36 may be concentrically surrounded by a shroud block assembly 40that is disposed within the turbine casing 38. The hot gas 30 rapidlyexpands as it flows through the turbine section 22. Thermal and/orkinetic energy is transferred from the hot gas 30 to each stage of therotor blades 36, thereby causing the shaft 24 to rotate and producemechanical work. The shaft 24 may be coupled to a load such as agenerator (not shown) so as to produce electricity. In addition, or inthe alternative, the shaft 24 may be used to drive the compressorsection 12 of the gas turbine.

FIG. 2 illustrates an enlarged side view of a portion of the turbinesection 22 including an exemplary rotor blade 36 and a portion of ashroud block assembly 40 according to various embodiments of the presentdisclosure. The turbine rotor blade or airfoil 36, extends from anaxially forward leading edge 44 to an axially aft trailing edge 46, andfrom a radially inward root 48 to a radially outer tip 42. The airfoil36 includes a platform 50 defining a radially inner boundary of a hotgas path. As shown in FIG. 2, the shroud block assembly 40 generallyextends in a radial direction 94 outward from the airfoil 36 between theturbine casing 38 (not shown) and a tip portion 42 of the rotor blade36. The shroud block assembly 40 generally includes mounting hardware 62for securing a plurality of shroud block segments 100 to the shroudblock assembly 40. The plurality of shroud block segments 100 may bearranged circumferentially 90 in an annular array around the rotorblades 36 within the turbine casing 38 (not shown).

Still referring to FIG. 2, each shroud block segment 100 may include aslash face 64 forming a circumferential interface with an adjacentshroud block segment 100. Stated otherwise, the plurality of shroudblock segments 100 may be arranged circumferentially 90, and the slashface 64 of each shroud block segment 100 may contact and/or be adjacentto the slash face 64 of an adjacent shroud block segment 100. Eachshroud block segment 100 may also include a shroud segment forward edge68, a shroud segment aft edge 66, and a shroud hot gas surface 76. Theshroud segment forward edge 68 and the shroud segment aft edge 66 arelocated on the axially forward and aft ends respectively of each shroudblock segment 100, while the shroud hot gas surface 76 is disposed at aradially inward surface of each shroud block segment 100 and forms theradially outer boundary of the turbine hot gas path. The platform 50 mayalso include a platform forward edge 72, a platform aft edge 70, and aplatform circumferential edge 74 disposed at the axially forward,axially aft and circumferential ends, respectively, of each platform 50.Coatings of the embodiments disclosed herein may be disposed at one ormore of the transition piece 52, the combustor liners 54, the firststage nozzle 56, the second stage nozzle 58, other turbine nozzles 34,flow-path ducts 60, the slash face 64, the shroud segment forward edge68, the shroud segment aft edge 66, the shroud hot gas surface 76, theplatform forward edge 72, the platform aft edge 70, the platformcircumferential edge 74, as well as on other seal surfaces of CMCcomponents.

FIGS. 3 and 4 summarize a series of test points that were undertaken toquantify various material properties of the coatings of the embodimentsdisclosed herein. In each case, a coating with a thickness between about0.01 inches and about 0.07 inches was applied to a crucible usingvarious methods. In one or more cases, a coating with a thicknessbetween about 0.02 inches and about 0.05 inches was used. Any cruciblewith suitable temperature resistance and surface qualities may be used.For example, the crucible may include one or more graphite crucibles,ceramic rods with rounded edges, machined ceramic surfaces, and/or othersuitable crucibles. The coating may be applied to the crucible via atape, mold, air spray, manually brushed on, applied as a powder andsubsequently crystallized, and/or via other suitable techniques. Inaddition, various modifications may be made to the application processto arrive at the desired coating thickness, density, etc. For example,the flow rate at which the coating is applied via an air sprayapplication process may be adjusted. In addition, the rate at which thecrucible is translated under the application equipment (and/or the rateat which the application equipment is moved over the crucible) may beadjusted. The height from which the coating is applied as well as thepressure at which the coating is applied may both also be adjusted.

FIG. 3 illustrates the processing temperatures 356 at which a number ofmaterial systems 352 including specific compositions 354 where tested,as well as the resulting coefficients of thermal expansion (CTE) 358 andsoftening points 360. Test points 302, 304, 306, and 308 were allperformed on a B-A-S material system including compositions with variousmass percentages of barium oxide (i.e., BaO and “B” in “B-A-S”),aluminum oxide (i.e., Al2O3 and “A” in “B-A-S”) and silicon dioxide(i.e., SiO2 and “S” in “B-A-S”). For example, test point 302 includes 33percent by mass silicon dioxide, 61.5 percent by mass barium oxide, and5.5 percent by mass aluminum oxide, while test point 304 includes 40.3percent by mass silicon dioxide, 51.3 percent by mass barium oxide, and8.4 percent by mass aluminum oxide, as indicated by the compositionpercentages listed in the composition column 354 of FIG. 3. Test point302, which was performed at a processing temperature 356 of 1300° C.,resulted in a CTE 358 of 12.4 (×10{circumflex over ( )}−6/(° C.)) and asoftening point 360 greater than 1200° C. Although test point 302resulted in a favorable softening point 360 greater than 1200° C., theCTE 358 was higher than desirable. Test point 304, which was alsoperformed at a processing temperature 356 of 1300° C. but on a differentB-A-S composition than test point 302, resulted in a CTE 358 of 8.9(×10{circumflex over ( )}−6/(° C.)) and a softening point 360 of 815° C.Thus, test point 304 resulted in neither a desired CTE 358 nor a desiredsoftening point 360. Both test points 306 and 308, which were performedon various B-A-S compositions at 1400° C. and 1300° C. respectivelyresulted in favorable CTE 358 (i.e., equal to or below about 8.0(×10{circumflex over ( )}−6/(° C.))), but lower than desirable softeningpoints 360. A CTE 358 was not quantified for test point 308.

Referring still to FIG. 3, test points 310, 312, and 314 were performedon a B-A-S/SiC material system including compositions by mass of B-A-Sand SiC (silicon carbide) as shown in the composition column 354. TheB-A-S material system that was used in each of test points 310, 312, and314 included 66.7 percent by mass silicon dioxide, 20 percent by massbarium oxide, and 13.3 percent by mass aluminum oxide (i.e., the sameB-A-S material system as test point 306). Test point 310 was performedon a composition of 98% B-A-S and 2% SiC while test point 312 wasperformed on a coating with a composition of 80% B-A-S and 20% SiC. Eachof test points 310, 312, and 314 were performed at a processingtemperature 356 of 1400° C. and resulted in favorable CTE 358 (i.e.,equal to or below about 8.0 (×10{circumflex over ( )}−6/(° C.))), butlower than desirable softening points 360. In addition, test points 312and 314 resulted in porous coatings.

Still referring to FIG. 3, test points 316, 318, and 320 were eachperformed on a S-A-Y material system including 54 percent by masssilicon dioxide (i.e., SiO2 and “S” in “S-A-Y”), 18.07 percent by massaluminum oxide (i.e., Al2O3 and “A” in “S-A-Y”), and 27.93 percent bymass yttrium oxide (i.e., Y2O3 and “Y” in “S-A-Y”). Test points 316,318, and 320 were performed at processing temperatures 356 of 1300° C.,1350° C., and 1400° C., respectively and each resulted in favorable CTE358 (i.e., equal to or below about 8.0 (×10{circumflex over ( )}−6/(°C.))) and favorable softening points 360 (i.e., greater than 1200° C.).However, test point 316 resulted in a porous coating.

Referring still to FIG. 3, test points 322, 324, 326, 328, 330, and 332were each performed on coating compositions include a mixture of S-A-Yand B2O3 material systems. The S-A-Y material system of each of testpoints 322, 324, 326, 328, 330, and 332 includes 54 percent by masssilicon dioxide (i.e., SiO2 and “S” in “SAY”), 18.07 percent by massaluminum oxide (i.e., Al2O3 and “A” in “S-A-Y”), and 27.93 percent bymass yttrium oxide (i.e., Y2O3 and “Y” in “S-A-Y”), similar to the S-A-Ymaterial system of test points 316, 318, and 320. Each of test points322, 324, 326, 328, 330, and 332 also include boron oxide (i.e., B2O3).For example, test point 322 includes 97 percent by mass S-A-Y materialand 3 percent by mass boron oxide while test points 324, 326, and 328each include 95 percent by mass S-A-Y material and 5 percent by massboron oxide. Each of test points 322, 324, 326, 328, 330, and 332resulted in a favorable CTE 358 (i.e., equal to or below about 8.0(×10{circumflex over ( )}−6/(° C.))), but only test points 322, 324,326, and 328 resulted in favorable softening points 360 (i.e., greaterthan 1200° C.). In addition, test points 322, 324, and 326 all resultedin porous coatings.

Still referring to FIG. 3, test points 334, 336, 338, 340, 342, 344,346, 348, and 350 were each performed on coating compositions includinga mixture of S-A-Y and B-A-S material systems, in different ratios andat different processing temperatures 356. Each of test points 334, 336,338, 340, 342, 344, 346, 348, and 350 include a S-A-Y material systemincluding 54 percent by mass silicon dioxide (i.e., SiO2 and “S” in“S-A-Y”), 18.07 percent by mass aluminum oxide (i.e., Al2O3 and “A” in“SAY”), and 27.93 percent by mass yttrium oxide (i.e., Y2O3 and “Y” in“S-A-Y”), similar to the S-A-Y material system of test points 316-332.Each of test points 334, 336, 338, 340, 342, 344, 346, 348, and 350include a B-A-S material system including 66.7 percent by mass silicondioxide, 20 percent by mass barium oxide, and 13.3 percent by massaluminum oxide (i.e., the same B-A-S material system as test points 306,310, 312, and 314). Each of test points 334, 336, 338, 340, 342, 344,346, 348, and 350 resulted in a favorable CTE 358 (i.e., equal to orbelow about 8.0 (×10{circumflex over ( )}−6/(° C.))). Test points 334,336, 338, 340, 342, 346, and 348 (i.e., all but test points 344 and 350)resulted in favorable softening points 360 (i.e., greater than 1200°C.). In addition, test points 334, 336 and 340 resulted in porouscoatings.

FIG. 4 illustrates a summary of each of the non-porous coatings thatinclude favorable CTE 358 and softening temperatures 360 from FIG. 3. Assuch, the coatings resulting from test points 318, 320, 328, 338, 342,346, and 348 are summarized included in FIG. 4. Coatings resulting fromthese seven test points were further studied to understand surfaceroughness. The coatings resulting from test points 318, 320, 328, 338,and 346 were found to have rough surfaces while the coatings resultingfrom test points 342 and 348 were found to have smooth surfaces. Thesurface finish was quantified in terms of roughness average, Ra, withunits in micro-inches (inches×10{circumflex over ( )}−6). RoughnessAverage, Ra is calculated as the average of a surface's measuredmicroscopic peaks and valleys, Rough surfaces may include surfaces witha roughness average of above about 150 Ra, or from about 150 Ra to about250 Ra. Smooth surfaces may include surfaces with a roughness average ofbelow about 150 Ra, or below about 100 Ra. In other embodiments, smoothsurfaces may include surfaces with a roughness average between about 10Ra and about 80 Ra. In other embodiments, smooth surfaces may includesurfaces with a roughness average between about 15 Ra and about 60 Ra.In other embodiments, smooth surfaces may include surfaces with aroughness average between about 20 Ra and about 50 Ra. In otherembodiments, smooth surfaces may include surfaces with a roughnessaverage between about 30 Ra and about 40 Ra. Smooth surfaces may includesurfaces with a roughness average between about 10 Ra and about 150 Ra,and all sub-ranges therebetween.

The compositions of the material systems included in FIGS. 3 and 4 mayvary around the exact constituents shown. For example, the S-A-Ycomposition including approximately 54% S, 18% A and 28% Y (i.e., testpoints 316, 318, and 320) may include tolerance bands of 2%, 3%, andeven 5% around each of the constituents. For example, the S-A-Ycomposition may include from about 49% to about 59% S, from about 13% toabout 23% A, and from about 23% to about 33% Y, assuming 5% tolerancebands. Similarly, the S-A-Y composition may include from about 51% toabout 57% S, from about 15% to about 21% A, and from about 25% to about31% Y, assuming 3% tolerance bands. Similarly, the S-A-Y composition mayinclude from about 52% to about 56% S, from about 16% to about 20% A,and from about 26% to about 30% Y, assuming 2% tolerance bands.Therefore, for the test points that include a mix of S-A-Y and B-A-Smaterial systems (i.e., test points 334-350) which includes as little as70% of S-A-Y material system in the overall composition, the overallcomposition may include as little as 70% of the lower limits of each ofthe ranges resulting from 5% tolerance bands. For example, theS-A-Y:B-A-S material systems in a 70:30 ratio may include as little asabout 34.3 percent by mass S (i.e. silicon dioxide), about 9.1 percentby mass aluminum oxide, and about 16.1 percent by mass yttrium oxide,when accounting for the constituents of the S-A-Y material system. Whenalso accounting for the B-A-S material system, in a 90:10 SAY:B-A-Sratio, the silicon dioxide may increase from about 34.3 percent to about40.5 percent (i.e., an addition of 6.2% of the overall composition dueto the silicon dioxide in the B-A-S material system, while the aluminumoxide may increase from about 9.1 to about 9.9 percent (i.e., anaddition of about 0.8% of the overall composition due to the aluminumoxide in the B-A-S material system).

The B-A-S material composition of test point 306 was repeated again fortest points 310-314 and 334-350 due to the low CTE 358. This materialcomposition included 66.7% S, 20% B, and 13.3% A. When surrounded by 5%tolerance bands, these percentages range from about 62% to about 72% S,about 15% to about 25% B, and about 8% to about 18% A. When surroundedby 3% tolerance bands, these percentages range from about 64% to about70% S, about 17% to about 23% B, and about 10% to about 16% A. Whensurrounded by 2% tolerance bands, these percentages range from about 65%to about 79% S, about 18% to about 22% B, and about 11% to about 15% A.

FIG. 5 illustrates a method 500 of forming CMC components coating withcoatings according to the embodiments disclosed herein. At step 502, themethod includes providing a CMC component. The CMC component may includea finished CMC component, a CMC component that is in a green state,and/or a CMC component that has been solidified but that may require oneor more post-processing steps such as heat treat and/or machining. Atstep 504, the method may include performing an EDM (electro dischargemachining) process on the CMC component. At step 506, the method 500 mayinclude performing a grinding process on the CMC component. Steps 504and 506 may be performed in any order. In some embodiments, both steps504 and 506 are performed. In some embodiments, neither step 504 norstep 506 is performed. In some embodiments, only one of steps 504 and506 is performed. At step 508, the method 500 may include depositing acoating on the CMC component. The coating may be deposited on the CMCcomponent via spray, plasma spray, and/or air spray, tape, manualbrushing, and/or via powder which may be subsequently crystallizedand/or reacted onto the CMC component. At step 510, the method 500 mayinclude heat treating the coated CMC component. At step 512, the method500 may include grinding the coated CMC component. At step 514, themethod 500 may include honing the coated CMC component. At step 516, themethod 500 may include extrude honing the coated CMC component. At step518, the method 500 may include polishing the coated CMC component. Insome embodiments, other steps may be performed. In some embodiments, oneor more of steps 502 through 518 may be omitted. In some embodiments,one or more of steps 502 through 518 may be performed in a differentorder that what is illustrated in FIG. 5.

Coatings and/or sealants of the embodiments disclosed herein may bedisposed at each of the locations discussed above in which a CMC surfaceand/or component defines a flow-path boundary and/or defines aninterfacing surface with an adjacent component (i.e., thereby forming aseal). Stated otherwise, it may be desirable to dispose the coating ofthe present embodiments on any CMC boundary and/or interfacing surfaceon which an enhanced seal and/or improved surface finish is desired. Thesurfaces on which the coatings may be disposed may include one or moreof the transition piece 52, the combustor liners 54, the first stagenozzle 56, the second stage nozzle 58, other turbine nozzles 34,flow-path ducts 60, one or more slash faces 64, one or more shroudsegment forward edges 68, one or more shroud segment aft edges 66, oneor more shroud hot gas surfaces 76, a platform forward edge 72, aplatform aft edge 70, a platform circumferential edge 74, as well as onother seal surfaces of CMC components.

Each of the coatings according to the embodiments disclosed herein mayinclude other oxides other than those listed above. For example, boronoxide may include boron dioxide, boron trioxide, boron monoxide, and/orboron suboxide. Similarly, aluminum oxide may include aluminum (I)oxide, aluminum (II) oxide, and/or aluminum (III) oxide. Each of thecoatings according to the embodiments disclosed herein may include aglassy surface that remains smooth in operation. By contrast, otherconventional EBCs (environmental barrier coating) and/or TBCs (thermalbarrier coatings) may be brittle and may chip away over time, whenexposed to gas turbine internal operating temperatures. Each of thecoatings according to the embodiments disclosed herein may include a CTE(coefficient of thermal expansion) that is substantially similar to thatof the CMC substrate on which they are disposed. For example, in someembodiments, the CTE of the coatings disclosed herein are within about50% of the CTE of the CMC substrate on which they are disposed. In otherembodiments, the CTE of the coatings disclosed herein are within about20% of the CTE of the CMC substrate on which they are disposed. In otherembodiments, the CTE of the coatings disclosed herein are within about10% of the CTE of the CMC substrate on which they are disposed. In otherembodiments, both the coatings disclosed herein as well as the CMCsubstrate on which they are disposed include a CTE that is equal to orless than about 8.0 (×10{circumflex over ( )}−6/(° C.)). Withconventional coatings, undulations may form on seal surfaces which mayallow undesired leakages to flow past. By including a CTE thatapproximately matches the CMC substrate on which they are disposed, byincluding a high softening temperature, and by remaining smooth whenexposed to internal gas turbine operating temperatures, the coatingsdisclosed herein may provide a smooth sealing surface substantially freefrom undulations, thereby resulting in enhanced sealing. Statedotherwise, the coatings disclosed herein may prevent mating and/orinterfacing CMC seal surfaces from developing undulations and/or gaps,which may lead to increases in undesired leakage flows.

Although specific features of various embodiments of the presentdisclosure may be shown in some drawings and not in others, this is forconvenience only. In accordance with the principles of the presentdisclosure, any feature of a drawing may be referenced and/or claimed incombination with any feature of any other drawing.

This written description uses examples to disclose the embodiments ofthe present disclosure, including the best mode, and also to enable anyperson skilled in the art to practice the disclosure, including makingand using any devices or systems and performing any incorporatedmethods. The patentable scope of the embodiments described herein isdefined by the claims, and may include other examples that occur tothose skilled in the art. Such other examples are intended to be withinthe scope of the claims if they have structural elements that do notdiffer from the literal language of the claims, or if they includeequivalent structural elements with insubstantial differences from theliteral language of the claims.

What is claimed is:
 1. A material composition, comprising: at least 34.9percent by mass silicon dioxide; at least 9.1 percent by mass aluminumoxide; at least 16.1 percent by mass yttrium oxide; and a B-A-S materialcomposition comprising barium oxide, aluminum oxide, and silicondioxide.
 2. The material composition of claim 1, further comprising: atleast 3 percent by mass boron oxide.
 3. The material composition ofclaim 2, further comprising: at least 5 percent by mass boron oxide. 4.The material composition of claim 3, further comprising: at least 41.7percent by mass silicon dioxide; at least 11.1 percent by mass aluminumoxide; and at least 19.6 percent by mass yttrium oxide.
 5. The materialcomposition of claim 1, wherein the B-A-S material compositioncomprises: from about 15 percent to about 25 percent by mass bariumoxide; from about 8 percent to about 18 percent by mass aluminum oxide;and from about 62 percent to about 72 percent by mass silicon dioxide.6. The material composition of claim 1, wherein the B-A-S materialcomposition comprises at least about 10% of a total composition of thematerial composition.
 7. The material composition of claim 1, whereinthe B-A-S material composition comprises between about 10% and about 30%of the total composition of the material composition.
 8. The materialcomposition of claim 1, further comprising: from about 49 percent toabout 59 percent by mass silicon dioxide; from about 13 percent to about23 percent by mass aluminum oxide; and from about 23 percent to about 33percent by mass yttrium oxide.
 9. The material composition of claim 1,further comprising: from about 51 percent to about 57 percent by masssilicon dioxide; from about 15 percent to about 21 percent by massaluminum oxide; and from about 25 percent to about 31 percent by massyttrium oxide.
 10. The material composition of claim 1, furthercomprising: from about 52 percent to about 56 percent by mass silicondioxide; from about 16 percent to about 20 percent by mass aluminumoxide; and from about 26 percent to about 30 percent by mass yttriumoxide.
 11. The material composition of claim 1, further comprising acoating of the material composition.
 12. The material composition ofclaim 11, further comprising a ceramic matrix composites (CMC)substrate, wherein the coating is disposed over the CMC substrate. 13.The material composition of claim 12, wherein the CMC substrate isporous, and the coating is non-porous.
 14. The material composition ofclaim 1, further comprising a component of a gas turbine engine havingthe material composition.
 15. The material composition of claim 1,wherein the material composition comprises a smooth surface having aroughness of less than about 150 Ra.
 16. The material composition ofclaim 1, wherein the material composition comprises a smooth surfacehaving a roughness of between about 10 Ra and about 80 Ra.
 17. Acoating, comprising: at least 34.9 percent by mass silicon dioxide; atleast 9.1 percent by mass aluminum oxide; and at least 16.1 percent bymass yttrium oxide, wherein the coating comprises a smooth surfacehaving a roughness of less than about 150 Ra.
 18. The coating of claim17, wherein the coating comprises a coefficient of thermal expansion(CTE) equal to or less than about 8.0 (×10{circumflex over ( )}−6/(°C.)), wherein the coating comprises a softening point of greater than1200° C., and the coating is non-porous.
 19. The coating of claim 17,wherein the coating is disposed over a ceramic matrix composites (CMC)substrate.
 20. The coating of claim 17, wherein the coating furthercomprises a B-A-S material composition, wherein the B-A-S materialcomposition comprises barium oxide, aluminum oxide, and silicon dioxide.21. A system, comprising: a component of a gas turbine engine, whereinthe component comprises a ceramic matrix composites (CMC) substrate; anda coating disposed over the CMC substrate, wherein the coating has amaterial composition, comprising: at least 34.9 percent by mass silicondioxide; at least 9.1 percent by mass aluminum oxide; and at least 16.1percent by mass yttrium oxide, wherein the coating comprises a smoothsurface having a roughness of less than about 150 Ra, and the coatingdefines a flow-path boundary and/or an interfacing surface in the gasturbine engine.
 22. The system of claim 21, wherein the materialcomposition of the coating further comprises a B-A-S materialcomposition, wherein the B-A-S material composition comprises bariumoxide, aluminum oxide, and silicon dioxide.